High-power ion propulsion systems have been shown to be capable of providing substantial benefits for the exploration of space. Ion engines have been under development for nearly thirty years and have long since achieved performance levels (thrust, specific impulse, and efficiency) which are attractive for planetary missions. The power-limited, low-thrust nature of ion propulsion, however, results in the requirement for very long engine burn times to produce the desired spacecraft velocity change. Engine burn times of 10,000 to 15,000 hours (201/2 months) are required for typical deep space missions of interest. Demonstrating useful engine lifetimes of this magnitude has historically proved to be an intractable problem, and yet such a demonstration is believed to be absolutely essential before this technology can be used on a spacecraft.
Considerable useful background material is contained in the following papers:
(1) James, E. L. and Bechtel, R. T., "Results of the Mission Profile Life Test First Test Segment: Thruster J1," AIAA Paper No. 81-0716, April 1981; PA1 (2) Bechtel, R. T., Trump, G. E., and James, E. J., "Results of the Mission Profile Life Test," AIAA Paper No. 82-1905, November 1982; PA1 (3) Collett, C., et. al., "Thruster Endurance Test," NASA CR-135022, May 1976; PA1 (4) Patterson, M. J. and Verhey, T. R., "5 kW Xenon Ion Thruster Lifetest," AIAA Paper No. 90-2543, July 1990; and PA1 (5) Rawlin, V. K. and Hawkins, C. E., "Increased Capabilities of the 30-cm Diameter Hg Ion Thruster," NASA TM-79142, May 1979.
A substantial effort has been expended to demonstrate adequate engine life, as represented by numerous life tests of the 30-cm diameter mercury ion engine in the 1970s and early 1980s. Not one of these tests, however, successfully demonstrated a 10,000- to 15,000-hour useful life at full-power operation. By far the most successful test was the 4,200-hour full-power test of the J-Series engine designated J1, as reported in the paper by James and Bechtel listed as (1) above. Other significant tests included a 5,000-hour test of the J5 thruster at one-quarter power described in the paper listed as (2) above, and a 10,000-hour test of a 700 series thruster described in the paper by Collett et al. listed as (3) above. The 10,000-hour test is considered significant only in that a 30-cm diameter thruster was operated for 10,000 hours. The thruster itself was nearly completely destroyed from internal sputter erosion by the end of the test, so that the test could not be considered a successful demonstration of a 10,000-hour engine life.
These and most other ion engine endurance tests have been performed using mercury propellant. Very low tank pressures (typically less 3.times.10.sup.-6 Torr) could be maintained during engine testing by cryopumping the mercury exhaust on surfaces at liquid nitrogen temperature in the vacuum system. Cryopumping results in very high effective pumping speeds if large surface areas are used, resulting in low background pressures. A major cost of mercury ion engine life testing turned out to be the cost of the liquid nitrogen required.
The switch from mercury to xenon propellants for ion engines made the already difficult life testing problem almost impossible. Xenon gas cannot be effectively cryopumped at liquid nitrogen temperatures. Furthermore, the heavy atomic mass of xenon greatly reduces the effective pumping speed of large oil-diffusion pumps. Finally, the 5-kW xenon thruster under development at NASA Lewis Research Center operates at a propellant flow rate that is 60% higher than the 2.7-kW J-Series mercury engine. The combination of these factors makes xenon engine life testing at low tank pressures very difficult and expensive.
Very low background pressures are required for ion engine life testing in order to minimize charge exchange erosion of the accelerator grid. The accelerator grid is the downstream electrode of the ion accelerator system, as shown in FIG. 1, and is typically biased several hundred volts negative of the beam plasma potential to shield the positive high-voltage engine from electrons produced by the neutralizer cathode. Charge-exchange ions created downstream of the engine, however, are accelerated into the negative accelerator grid and cause sputter erosion.
The production rate of charge-exchange ions is a strong function of the background pressure, as indicated in FIG. 2. The increase in accelerator grid current evident in FIG. 2 results from the increase in charge-exchange ion production with tank pressure. The datum point labeled "LeRC 890 HR LIFE TEST" represents the accelerator grid current and tank pressure at which NASA Lewis Research Center performed their 5-kW ion engine endurance test, described by M. J. Patterson and T. R. Verhey in the reference listed as (4) above. The solid line in FIG. 2 is a curve fitted to data taken at NASA Lewis Research Center and indicates that the accelerator grid current asymptotically approaches a minimum value of 11.2 mA at zero background pressure.
The 5-kW endurance test described in the paper by Patterson and Verhey listed as (4) above was performed with an accelerator current of 18N2 mA, which is approximately 60% higher than the zero pressure current. At this current level and an accelerator grid voltage of -300 V, erosion resulted in holes completely through the accelerator grid webbing in less than 890 hours. The solid line in FIG. 2 indicates that the accelerator current can be reduced to 12 mA at a background pressure of 10.sup.-6 Torr.
However, to achieve a background pressure of 10.sup.-6 Torr during full-power operation of the 5-kW engine requires a pumping speed of approximately 700,000 liters/s. To put this number in perspective, the large tank 5 at NASA Lewis Research Center with two He cryopanels will have a pumping speed on Xe of 250,000 liters/s. In addition, the Lewis Research Center Space Power Facility (SPF), with 36 48" oil diffusion pumps has a xenon pumping speed of 1.4.times.10.sup.6 liters/s. However, it has been estimated that the cost of operating this facility may be prohibitively expensive for long-duration engine testing.
Finally, the datum point labeled "JPL 2-GRID" on FIG. 2 indicates the accelerator grid current measured in the NASA Jet Propulsion Laboratory 8'.times.15' electric propulsion test facility at the minimum achievable tank pressure during engine operation at 5 kW. This accelerator grid current is approximately a factor of 3 higher than that for the Lewis Research Center 890-hour endurance test, suggesting that erosion holes would be created through the grid webbing in less than 300 hours if an endurance test were attempted in the NASA Jet Propulsion Laboratory facility.
With this background information, it is abundantly clear that life testing a 5-kW xenon ion engine is an extremely expensive undertaking. One approach to the problem is to defer the cost of the engine life demonstration to the first flight project which will use the technology. However, attempts to get a flight project to accept the cost and scheduling impacts associated with long-duration life testing have historically been unsuccessful, and there seems to be little indication that this situation will change. Thus the substantial benefits of ion propulsion are trapped within a programmatic catch-22, i.e. it is too expensive to life-test an engine without the financial resources of a flight project, and a flight project is unlikely to provide these resources for a new engine technology that has not already demonstrated that it has the required engine life.